This invention relates to an electric power processing device, and more particularly, to an integrated aircraft power conditioning and control unit that interfaces generation equipment with various load equipment utilizing independent voltages levels and frequencies.
Many industries can benefit from lightweight power conditioning systems that are also flexible in providing a variety of voltages of different magnitudes and frequencies. One such industry is the aviation industry. For example, advances in unmanned aircraft are necessitating new electric power system architectures that allows the unmanned aircraft to be autonomous while minimizing the size and weight of the aircraft. The related art equipment currently being used to support such prototype aircraft is not optimal with respect to the size and weight of the electric power system. A flexible power conversion system will allow power to flow between various “flavors” of electricity (i.e., between AC and DC and between low voltage and high voltage) to achieve various modes of operation onboard modern aircraft. The various modes of operation may include engine start, ground support, normal flight and emergency operation.
FIG. 1 illustrates a related art electric power system for an aircraft. The electric power system includes generator 10, generator control unit (GCU) 15, External Power Connection (EPC) DC ground cart interface 40, high voltage battery 20, low voltage battery 30, DC-DC converter 50 and inverter 60. A typical aircraft may have two electrical power systems similar to that illustrated in FIG. 1.
Generator 10 includes a wound field synchronous machine (WFSM) 12 that is configured to be used as a generator. Generator 10 also includes a permanent magnet generator (PMG) 13 that supplies control power to GCU 15. WFSM 12 and PMG 13 are both mounted on a shaft from engine gear box 5. The output of generator 10 forms a high voltage DC bus 25 by rectifying the output of WFSM 12 using rectifier 11. GCU 15 controls the excitation voltage of WFSM 12 to maintain a desired DC voltage at the output of generator 10. In this configuration, high voltage DC bus 25 is the source for all the electrical power for the aircraft, and high voltage DC bus 25 may, for example, have a magnitude of 270 volts.
Connected to high voltage DC bus 25 is high voltage battery 20. During normal operation, the charge on high voltage battery 20 is maintained by generator 10 via high voltage DC bus 25. A battery charger and disconnect switches (both features not shown) may be connected between the high voltage battery 20 and high voltage DC bus 25. When generator 10 is not available or if the power from generator 10 is insufficient, high voltage battery 20 provides power to high voltage DC bus 25 to operate the equipment.
The input power to DC-DC converter 50 is provided by high voltage DC bus 25, and the output of DC-DC converter 50 forms a low voltage DC bus 35 that supplies control power to the system avionics. During normal flight operation, the charge on low voltage battery 30 is maintained by DC-DC converter 50 via low voltage DC bus 35. A battery charger and disconnect switches (both features not shown) may be connected between low voltage battery 30 and the low voltage DC bus. If DC-DC converter 50 is not operational or if the power from DC-DC converter 50 is insufficient, low voltage battery 30 will provide power to low voltage DC bus 35. The magnitude of low voltage DC bus 35 may be, for example, 28 volts.
Inverter 60 provides power to legacy equipment that run on AC power. Inverter 60 gets its supply from high voltage DC bus 25 and converts the DC power to AC power at, for example, 115 volts, 400 Hz.
EPC DC ground cart interface 40 is connected to high voltage DC bus 25 and allows for external power to supply the aircraft when the aircraft is on the ground (shore power). For example, EPC DC ground cart 41 provides power to the 270 VDC equipment and the 115 volt, 400 Hz equipment via inverter 60.
Alternatively, an EPC AC ground cart (not shown in FIG. 1) may be connected to the legacy AC bus to provide power to the legacy equipment when the aircraft is on the ground. However, the related art system described above does not allow the EPC AC ground cart to assist high voltage battery 20 in main engine startup (main engine startup circuitry is not shown in FIG. 1). Adding additional circuitry to permit the EPC AC ground cart to assist in the main engine start will add additional weight and complexity to the power conditioning system of the aircraft because the AC power will have to be converted to DC.
In addition, the modular design of the related art control system also adds to the complexity and weight of the power conditioning system. FIG. 2 illustrates a control block diagram for the related art modular control system. The functional control and fault management functions used in making strategic decisions in the power conditioning unit are performed at the “highest” level in the vehicle control module 70.
At an “intermediate” level, the bus power control unit module (BPCU) 80 monitors the EPC controls and the left/right half bus protection and controls, i.e., the two generator systems of the aircraft.
At a “low” level, individual control circuit modules such as the GCU module 90 and the Inverter module 95 perform the “detailed” functions necessary to control the respective equipment. For example, the GCU module 90 has a voltage regulation block to maintain the voltage from generator 10 at a preset value and a protection and breaker control block to protect generator 10 from damage. Inverter module 95 has an engine start inverter control block that controls the start of the main engines (not shown in FIG. 1) and a legacy load control block that maintains the AC voltage on the legacy AC bus.
The separate nature of the control modules shown in FIG. 2 add to wiring mass, complexity and interface issues in the related art power conditioning systems. In contrast, the present invention provides a flexible, integrated power conditioning system that permits EPC AC or DC ground carts to supplement the high voltage batteries in main engine startup. By integrating the control modules of various power conditioning unit components into, for example, a single control module, the weight of the system can be lessened because the wires connecting the various individual control modules will be eliminated. In addition, the overall complexity and interface issues will be lessened. Moreover, the flexible nature of the power conditioning system will allow the onboard battery to be sized appropriately, thereby, allowing for an additional reduction in weight.